High-speed-launch ramjet booster

ABSTRACT

A high-speed-launch ramjet boost (HSLRB) engine includes a combustion system for igniting fuel pumped by a fuel pump from a fuel tank, where the combustion system includes an igniter, fuel injectors and frame holders. An inlet provides a pathway for air to flow toward the fuel injectors. A variable geometry (VG) nozzle having a nozzle actuator is included for exhausting exhaust gas from combustion of the fuel by the combustion system. A processor is coupled to receive sensing signals from at least one of a pressure sensor and a temperature sensor during flight, wherein the processor provides control signals to the nozzle actuator for dynamically controlling an aperture size of the VG nozzle.

CROSS REFERENCE TO RELATED APPLICATIONS

This application claims the benefit of Provisional Application Ser. No.61/723,906 entitled “HIGH-SPEED LAUNCH RAMJET BOOSTER”, filed Nov. 8,2012, which is herein incorporated by reference in its entirety.

FIELD

Disclosed embodiments relate to ramjet engines and ramjet-powered boostvehicles therefrom.

BACKGROUND

A ramjet engine (or stovepipe jet, or athodyd) is a form ofair-breathing jet engine which uses the forward motion of the engine tocompress incoming air which is fed via an inlet, without the need for arotary air compressor. Ramjets have historically been used as cruiseengines to power high-speed (typically Mach 2.5-3.5) missiles.

Thrust for the ramjet is produced by passing hot exhaust generated fromthe combustion of a fuel through a jet nozzle. The nozzle acceleratesthe flow, and the reaction to this acceleration produces thrust. Tomaintain the exhaust flow through the nozzle, the combustion must occurat a pressure higher than the pressure at the nozzle exit. In a ramjet,this needed high relative pressure is produced by “ramming” external airinto the combustor using the forward speed of the vehicle. Conventionalramjets have a fixed geometry (FG) inlet and a FG nozzle.

The minimum operating speed for free flight is set by the particularramjet design. The inlet and nozzle design will determine the minimumoperating speed that will yield excess thrust (thrust minus drag) forseparation and acceleration. When launched from a subsonic aircraft,ramjet-powered vehicles generally require a separate booster motor orvehicle to accelerate the ramjet to at least its minimum operating speed(typically Mach 2+) before lighting. This booster is typically a solidrocket, which significantly increases the size of the ramjetengine/vehicle.

SUMMARY

This Summary is provided to introduce a brief selection of disclosedconcepts in a simplified form that are further described below in theDetailed Description including the drawings provided. This Summary isnot intended to limit the claimed subject matter's scope.

Disclosed embodiments include ramjet engines and ramjet-powered boostvehicles therefrom. Disclosed ramjet engines are generally referred toas high-speed-launch ramjet boost (HSLRB) engines, and do not require aconventional booster, such as a conventional solid rocket booster. Inoperation, the HSLRB engine is adapted to be launched from a high-speedaircraft, be ignited at a supersonic speed while still being attached tothe aircraft, and generate enough excess thrust to enable launch fromthe aircraft at Mach 2+.

The HSLRB engine includes a variable geometry (VG) nozzle, and either afixed geometry (FG) inlet or an optional VG inlet. The VG nozzle managesthe inlet terminal shock and through actuation by its actuator providesnozzle aperture expansion throughout the large (Mach 3.5+) speed range.The VG inlet can be incorporated to provide more excess thrust at thelow end of the ramjet's speed range, if desired, such as to supportlaunch from a particular aircraft.

Disclosed HSLRB engines can be used as a first stage for air-launchedmicrosatellites, with an additional rocket-powered stage(s) used fororbital insertion. By disclosed embodiments employing a high-speedsupersonic aircraft to carry a disclosed HSLRB engine to ≧Mach 2+ priorto ramjet launch, the size of the HSLRB engine can be substantiallyreduced by eliminating the need for a rocket booster (e.g., solid rocketbooster) required for conventional ramjets. Thus, compared with otherair-launch schemes, a HSLRB stage for microsatellite launch offersadvantages including a significant decrease in overall vehicle mass andsize, a smaller logistic footprint, decreased launch costs, and moreeasily adaptable mission profiles.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a simplified depiction of an example microsatellite launchvehicle having a disclosed HSLRB engine, showing major components of theHSLRB engine, according to an example embodiment.

FIG. 2A and 2B are simplified depictions illustrating alternateembodiments utilizing the HSLRB engine shown in FIG. 1, with FIG. 2Ashowing an example ballistic launch vehicle with a disclosed HSLRBengine, while FIG. 2B shows a depiction of a high-speed cruise vehiclewith a disclosed HSLRB engine, according to example embodiments.

DETAILED DESCRIPTION

Disclosed embodiments are described with reference to the attachedfigures, wherein like reference numerals are used throughout the figuresto designate similar or equivalent elements. The figures are not drawnto scale and they are provided merely to illustrate aspects disclosedherein. Several disclosed aspects are described below with reference toexample applications for illustration.

It should be understood that numerous specific details, relationships,and methods are set forth to provide a full understanding of theembodiments disclosed herein. One having ordinary skill in the relevantart, however, will readily recognize that the disclosed embodiments canbe practiced without one or more of the specific details or with othermethods. In other instances, well-known structures or operations are notshown in detail to avoid obscuring aspects disclosed herein. Disclosedembodiments are not limited by the illustrated ordering of acts orevents, as some acts may occur in different orders and/or concurrentlywith other acts or events. Furthermore, not all illustrated acts orevents are required to implement a methodology in accordance with thisDisclosure.

Notwithstanding that the numerical ranges and parameters setting forththe broad scope of this Disclosure are approximations, the numericalvalues set forth in the specific examples are reported as precisely aspossible. Any numerical value, however, inherently contains certainerrors necessarily resulting from the standard deviation found in theirrespective testing measurements. Moreover, all ranges disclosed hereinare to be understood to encompass any and all sub-ranges subsumedtherein. For example, a range of “less than 10” can include any and allsub-ranges between (and including) the minimum value of zero and themaximum value of 10, that is, any and all sub-ranges having a minimumvalue of equal to or greater than zero and a maximum value of equal toor less than 10, e.g., 1 to 5.

FIG. 1 is a simplified depiction of a microsatellite launch vehicle 100having a disclosed HSLRB engine 110, showing major components of theHSLRB engine 110. In addition to the microsatellite launch vehicle 100shown, as described below, the HSLRB engine 110 can also be used for aballistic launch vehicle, or a high-speed cruise vehicle (see FIG. 2Aand FIG. 2B, respectively, described below). The microsatellite launchstage presents a challenging set of system requirements, chief of whichis generally to minimize mass, and is described in some detailed below.Briefly, a small increase in microsatellite launch stage mass cansignificantly decrease the payload mass.

For the microsatellite launch vehicle 100 as shown in FIG. 1, the HSLRBengine 110 is located behind a separate orbital injection stage 130having at least one payload 135. The microsatellite launch vehicle 100includes a frame 101 having a front portion 101 a and an aft portion 101b which provides an outside framing structure for the orbital injectionstage 130 and the HSLRB engine 110. The HSLRB engine 110 together withthe frame 101, the fuel tank 127 and fuel pump 128 is referred to hereinas a “HSLRB stage”.

HSLRB engine 110 includes an inlet (or inlets), shown as an optional VGinlet 111 with an associated inlet actuator 111 a. VG inlet 111 suppliesmore thrust at lower speeds compared to a conventional fixed inlet, withits utility reduced at faster speeds. However, as noted above, HSLRBengine 110 may alternately include a conventional fixed inlet.

HSLRB engine 110 includes a process/controller 138 (e.g., digital signalprocessor) hereafter processor 138, which is coupled to receive sensingsignals from at least one of a pressure sensor and a temperature sensorduring flight, with both a pressure sensor 136 and temperature sensor137 being shown in FIG. 1. The processor 138 is programmed by disclosedalgorithms in memory 131 that based on the level of the sensing signalsprovide control signals to the inlet actuator 111 a for dynamicallycontrolling a geometry of the VG inlet 111, which is described in moredetail below.

HSLRB engine 110 also includes a VG nozzle (or nozzles) 121 having anassociated nozzle actuator 121 a. The processor 138 is programmed byprograms in memory 131 that based on the level of the sensing signals toprovides control signals to the nozzle actuator 121 a for dynamicallycontrolling an aperture size (or throat size) of the VG nozzle 121,again described in more detail below.

Although VG nozzle 121 will tend to increase the complexity and cost ofthe HSLRB engine, the resulting propulsion system having a VG nozzle 121has been recognized herein to provide a very wide Mach/altitude envelopeby enabling a broad range of efficient operating Mach and altitude. Adisclosed VG nozzle 121 provides high specific impulse (Isp) andthrottleability that can support multiple missions, where the Isp andlow-speed excess thrust may be enhanced by including a VG inlet 111.

As noted above, the dimensions of the VG inlet 111 and the size of theVG nozzle 121 are dynamically controlled by control signals provided byprocessor 138 running algorithms stored in memory 131 based on sensingsignals from a pressure sensor 136 and/or temperature sensor 137. Asknown in the art, airspeed is derived from a combination of dynamic andstatic pressure, while Mach is derived from pressure and temperature. Ananalog to digital converter (ADC) for digitizing the sensing signals anda low pass filter, although generally provided, are not shown in FIG. 1for simplicity. Typically, the size of the nozzle aperture of the VGnozzle 121 will be controlled to be largest at low Mach and smallest athigh Mach to maintain the inlet terminal shock within acceptable bounds,and to provide near-ideal nozzle expansion.

At low Mach, the aperture of the VG nozzle 121 will be controlled to beat least substantially open, and it will be controlled to be closed downas Mach and ram pressure increase. In general, the aperture of the VGnozzle 121 is as closed (small) as possible to maximize the pressure,while avoiding increasing the pressure too rapidly, which can push theterminal shock out of the inlet. If a VG inlet 111 is used, the geometryof the VG inlet 111 will be controlled to maintain shock-on-lipoperating conditions across the Mach range, thereby maximizing thrustand minimizing spillage drag.

HSLRB engine 110 includes a fuel tank 127 coupled to a fuel pump 128,such as an air-driven turbopump, to feed the fuel into fuel injectors,shown as combustion system 129 including fuel injectors, flameholdersand igniter. Combustion is initiated by an igniter, generally eitherelectrical or pyrotechnic, and is maintained by flameholders provided byfuel injectors in combustion system 129. The liquid fuel is generally ahydrocarbon, typically jet fuel or some similar formulation. Exhaust gasfrom the combustion of the fuel flows out through VG nozzle 121 toprovide thrust for the microsatellite launch vehicle 100. Otherconfigurations are also possible; e.g., a pressurized-gas system couldreplace the turbopump, or a solid fuel could be used with no tank, pumpor injectors.

The VG inlet 111 can include movable ramps, a translating plug, or someother mechanism to provide near-isentropic compression and maintainshock-on-lip conditions in the primary-speed range. The specific type orlocation of VG inlet(s) 111 is generally not important, and a variety ofdifferent configurations for VG inlet 111 can be selected to meetpackaging (i.e. placement of internal components) constraints. Forsimplicity, and to establish a conservative baseline, concept efficacywas assessed with a single, ventral external-compression inlet with VGhorizontal compression ramps. An inlet capture area of 0.3 to 0.7(approximately ½) the frontal area of the vehicle will generally providea sufficient excess thrust to support a launch Mach of 2.2 and a stagingMach of 5.5+. This capture area also yields a maximum exit diameter forthe VG nozzle 121 equal to the outside diameter of the vehicle so theHSLRB stage's Isp and thrust can be maximized across wide ranges of Machand altitude.

The VG inlet 111 can also be designed to operate with the inlet cover117 shown in FIG. 1. The inlet cover 117 can comprise a frangible coverthat is present only prior to HSLRB engine 110 ignition (shatteredbefore igniting). An inlet cover 117 minimizes vehicle drag andeliminates inlet buzz, but adds some complexity and possibly some weightto the design of the VG inlet 111. These disadvantages may be traded offagainst the drag/buzz advantages to determine whether a cover mode isincluded. This determination might be launch aircraft specific, sodifferent inlet designs might be employed for different launch aircraft.

No inlet cover may be used if the HSLRB engine is employed to help thelaunch aircraft accelerate to launch speed. In this embodiment, inlet“buzz” conditions would be avoided, or the duct constructed to be strongenough to survive a transient buzz. The added inlet drag due to lack ofan inlet cover would not generally be an issue because the HSLRB engine110 generally produces enough thrust to more than offset inlet drag.

As noted above, a frangible inlet cover can be used. A frangible inletcover will generally be shattered just prior to ignition, and the piecesingested into the VG inlet 111 and then expelled through the VG nozzle121. As noted above, cover mode is also an option: With a VG inlet 111,the compression ramps, compression cone/plug, or translating cowl, couldbe moved in a manner that blocks most, of not all, of the flow into theVG inlet 111. If inlet buzz and drag are concerns, and a VG inlet isused, a cover mode could be an appropriate design.

The VG nozzle 121 is generally more important to the HSLRB engine 110efficiency as compared to a turbojet or rocket because at lower Machnumbers, the ram pressure on the VG nozzle 121 is relatively low. The VGnozzle 121 is also generally important to maintaining critical inletperformance across a wide Mach range. Conventional boosted ramjets canemploy fixed geometry nozzles because the booster accelerates the ramjetto a cruise Mach where the nozzle pressure ratio is higher. The specificform of VG nozzle 121 is generally not important, and can be 2D, 3D, oreven fluidic. Efficiency, weight, complexity, cost, and packaging candrive the VG nozzle 121 type selection.

Although not shown, the HSLRB engine 110 will generally include aelectrical generator or other source of electrical power (e.g., battery)to provide electrical power where needed to power the igniter (at leastinitially), the processor 138, and the nozzle actuator 121 a and inletactuator 111 a if a VG inlet 111 is provided. Where an air-driventurbopump is used for the fuel pump 128, the same turbopump can drive agenerator. It is noted the VG nozzle 121 and VG inlet 111 can bepositioned using hydraulic or pneumatic instead of electrical actuators,where the controls for the hydraulics or pneumatics receive electricalpower.

The non-electrical components of the HSLRB engine 110 can be constructedof a high-temperature-resistant material (e.g., metal alloy), anddesigned as a hot structure (i.e. a structure where part of the primarystructure is not insulated from aerodynamic heating). This simplifiesthe design, although there might be a weight penalty. There is no needto use a hot structure for the orbital injection stage 130. Weight isgenerally not as important on the HSLRB engine 110 as on the orbitalinjection stage 130.

As noted above, an enabling aspect to operation of the HSLRB engine 110is a high-speed launch from a supersonic launch aircraft. Nominal launchspeed is about Mach 2.2, but this can be varied by at least about 0.2Mach depending on specific mission/payload requirements. In typicaloperation of the microsatellite launch vehicle 100, the HSLRB engine 110is started (ignited) under the launch aircraft where it is held at asupersonic speed. If the launch aircraft has sufficient excess thrust toaccelerate to launch conditions (e.g., ≧2.0 Mach) with no assistance,the HSLRB engine 110 can be started immediately before launch. If thelaunch aircraft needs assistance during acceleration, the HSLRB engine110 can be started at a lower Mach (e.g., ˜1.5 Mach).

Regarding a typical operation concept, if the HSLRB engine 110 is notneeded to help the launch aircraft accelerate, the HSLRB engine can bestarted and then launched almost immediately thereafter. The HSLRBengine 110 can be started with a very rich fuel mixture to assure easyignition, then the fuel control controlled by processor 138 or anotherprocessor can revert to a schedule that maximizes thrust-specific fuelconsumption. The HSLRB engine 110 can go through an automatedbuilt-in-test (BIT) cycle, to insure that all actuators (nozzle actuator121 a and optional inlet actuator 111 a) are working properly, then canrevert back to a stable idle after BIT is completed. Given a successfulBIT, a retaining bolt may be retracted in the launcher, leaving only ashear bolt to restrain the HSLRB engine 110. The HSLRB engine 110 canthen go to full throttle, and when excess thrust exceeds the shearstrength of the restraint bolt, the microsatellite launch vehicle 100would leave the rail.

If the launch pylon is plumbed, aircraft internal fuel can be used torun the HSLRB engine 110 during captive carry by either adding to jetfuel in the ramjet's fuel tank 127, or through a bypass line that wouldprovide jet fuel to the fuel pump 128. Use of a bypass line would allowfor captive carry operation on aircraft internal fuel while stillproviding a separate supply of fuel that is generally more compatiblewith higher temperatures associated with post-launch operation. The VGinlet 111 of the HSLRB engine 110 will generally maximize availablethrust during sub-launch-Mach operation, and the thrust will offset dragof the microsatellite launch vehicle 100 and will assist launch aircraftacceleration until Mach 2+ is achieved.

Following HSLRB engine 110 launch, the HSLRB engine 110 can accelerateto its maximum Mach. The maximum speed can generally approach Mach 6.0,but acceleration drops considerably above Mach 5.5, so higher speedstypically offer diminishing returns. When acceleration is completed, theHSLRB engine 110 can begin a climb to staging conditions. For example,staging can occur at a nominal altitude/speed of 110,000 ft./Mach 5.5,although staging conditions can vary with mission requirements. Notably,the ability of the HSLRB engine 110 to achieve thesehigh-speed/angle/altitude staging conditions allows the orbitalinjection stage 130 to employ a high-expansion-ratio rocket nozzle andto employ a simple gravity turn with minimal impact on drag/gravitylosses, usually expressed as a change in velocity (ΔV). The orbitalinjection stage 130 can then place the payload 135 into an ellipticalparking orbit, which can be circularized if desired.

FIG. 2A and FIG. 2B are simplified depictions that illustrate alternateembodiments of the HSLRB engine 110 configured as a ballistic launchvehicle 200 with HSLRB engine 110 and a high-speed cruise vehicle 250with HSLRB engine 110, respectively. For use of HSLRB engine 110 as aballistic launch vehicle 200, this embodiment is similar to themicrosatellite launch vehicle 100 except that the latter stage(s) has norequirement to achieve orbit. The HSLRB engine 110 is generally providedsufficient fuel for an extended range. The vehicle stage split can bealtered to provide cruise range on the HSLRB engine 110; and the payload135 can be increased above that possible to deliver to the orbitalinjection stage 130, and/or the configuration of the latter stage(s)could be substantially altered to meet mission needs. This embodimentcovers all multi-stage, non-orbital vehicles that employ disclosed HSLRBengine launch schemes.

For the high-speed cruise vehicle 250 shown in FIG. 2B, the HSLRB engine110 is generally a single-stage, long range, high-speed cruise vehicle.Depending on altitude and range requirements, wings, larger fins, or alifting-body/waverider fuselage might be employed. The payload 135 canbe substantially increased, and considerable additional ramjet fuel canbe carried, depending on mission needs. This embodiment covers allsingle-stage, high-speed cruise vehicles that employ disclosed HSLRBlaunch schemes.

Regarding features believed to be unique regarding operation of adisclosed HSLRB engine 110, one such feature is high-speed launch. Theuse of a high-speed (e.g., ≧Mach 2) launch aircraft instead of aconventional booster enables the HSLRB engine/stage to operate at highIsp across the range of approximately Mach 2 through Mach 5.5. A ramjethas approximately 3 to 4times the Isp of a rocket across this speedrange, thus significantly reducing the size and mass of the stagerequired to transit this speed range. The VG nozzle 121 (particularlywith VG inlet 111) enable a broad range of efficient operating Mach andaltitude.

The operation Mach/altitude are also believed to be unique for ramjets.Disclosed HSLRB engines are also believed to be unique in its use for anunmanned air-launched vehicle that can operate solely in thesupersonic/hypersonic regime. The use of a ramjet with high excessthrust allows a rocket-propelled orbital injection stage to be startedat high speed, high altitude and high angle. High speed is generallyimportant, and is the most important of these three factors. High angleallows the orbital injection stage to rapidly accelerate with a minimumof drag loss and gravity ΔV loss compared with horizontal staging.

A high flight path angle at staging of approximately 30 degrees allows agravity turn to be employed for much of the orbital injection profilewith a minimum increase in gravity and drag ΔV loss. High altitudereduces aerodynamic drag and allows a high-expansion-ratio nozzle to beemployed on the orbital injection stage rocket, maximizing the Isp ofthat engine, which helps maximize orbital payload.

An advantageous application for a disclosed HSLRB engine is as a stagein an air-launched microsatellite launch vehicle, such as microsatellitelaunch vehicle 100 shown in FIG. 1 described above. Another advantageousapplication is as a stage in a ballistic suborbital vehicle or weapon,analogous to ballistic launch vehicle 200 shown in FIG. 2A describedabove. Yet another advantageous application is for propulsion of ahigh-speed cruise vehicle for intelligence, surveillance andreconnaissance (ISR) or research, analogous to high-speed cruise vehicle250 shown in FIG. 2B described above.

Significant advantages of disclosed HSLRB engines include materiallyreducing the launch footprint and cost of microsatellite air-launch. Fora given payload, the HSLRB can reduce the size and mass of the satellitelaunch vehicle's first stage by a factor of approximately 4×. DisclosedHSLRB engines can also reduce the size and mass of a latter stage(s) byproviding higher speed/angle/altitude staging conditions. Thesereductions in size and mass increase the available launch platforms fromcustom dedicated assets, such as Virgin Galactic's White Knight 2, tothe USAF's entire fleet of F-15s and privately-operated Mach 2 fighters.Competing proposals to a 2,300 lb disclosed HSLRB engine generally weighfrom 10,000 to 25,000 lb., with approximately the same payload. Inaddition, disclosed HSLRB engines offer a smaller logistics footprint,lower launch costs, and increased mission flexibility.

While various disclosed embodiments have been described above, it shouldbe understood that they have been presented by way of example only, andnot as a limitation. Numerous changes to the disclosed embodiments canbe made in accordance with the Disclosure herein without departing fromthe spirit or scope of this Disclosure. Thus, the breadth and scope ofthis Disclosure should not be limited by any of the above-describedembodiments. Rather, the scope of this Disclosure should be defined inaccordance with the following claims and their equivalents.

Although disclosed embodiments have been illustrated and described withrespect to one or more implementations, equivalent alterations andmodifications will occur to others skilled in the art upon the readingand understanding of this specification and the annexed drawings. Whilea particular feature may have been disclosed with respect to only one ofseveral implementations, such a feature may be combined with one or moreother features of the other implementations as may be desired andadvantageous for any given or particular application.

The terminology used herein is for the purpose of describing particularembodiments only and is not intended to be limiting to this Disclosure.As used herein, the singular forms “a,” “an,” and “the” are intended toinclude the plural forms as well, unless the context clearly indicatesotherwise. Furthermore, to the extent that the terms “including,”“includes,” “having,” “has,” “with,” or variants thereof are used ineither the detailed description and/or the claims, such terms areintended to be inclusive in a manner similar to the term “comprising.”

Unless otherwise defined, all terms (including technical and scientificterms) used herein have the same meaning as commonly understood by oneof ordinary skill in the art to which this Disclosure belongs. It willbe further understood that terms, such as those defined in commonly-useddictionaries, should be interpreted as having a meaning that isconsistent with their meaning in the context of the relevant art andwill not be interpreted in an idealized or overly formal sense unlessexpressly so defined herein.

1. A high-speed-launch ramjet boost (HSLRB) engine, comprising: Acombustion system for igniting fuel pumped by a fuel pump from a fueltank, said combustion system comprising an igniter, fuel injectors andframe holders; at least one inlet providing an pathway for air to flowtoward said fuel injectors; a variable geometry (VG) nozzle having anozzle actuator for exhausting exhaust gas from combustion of said fuelby said combustion system, and a processor coupled to receive sensingsignals from at least one of a pressure sensor and a temperature sensorduring flight, wherein said processor provides control signals to saidnozzle actuator for dynamically controlling an aperture size of said VGnozzle.
 2. The HSLRB engine of claim 1, wherein said inlet comprises aVG inlet having an inlet actuator, and said processor provides controlsignals to said inlet actuator for dynamically controlling a geometry ofsaid VG inlet.
 3. The HSLRB engine of claim 2, wherein said VG inletincludes an inlet cover.
 4. A launch vehicle, comprising: ahigh-speed-launch ramjet boost (HSLRB) stage including: a frameincluding a front portion and an aft portion, and a fuel tank and fuelpump within said frame; a high-speed-launch ramjet boost (HSLRB) enginewithin said frame including: a combustion system for igniting fuelpumped by said fuel pump from said fuel tank, said combustion systemcomprising an igniter, fuel injectors and frame holders; at least oneinlet providing an pathway for air to flow within said frame toward saidfuel injectors; a variable geometry (VG) nozzle having a nozzle actuatorat said aft portion for exhausting exhaust gas from combustion of saidfuel by said combustion system, and a processor coupled to receivesensing signals from at least one of a pressure sensor and a temperaturesensor during flight, wherein said processor provides control signals tosaid nozzle actuator for dynamically controlling a aperture size of saidVG nozzle, and at least one other stage including a payload attached tosaid aft portion of said HSLRB stage.
 5. The launch vehicle of claim 4,wherein said inlet comprises a VG inlet having an inlet actuator, andsaid processor provides control signals to said inlet actuator fordynamically controlling a geometry of said inlet.
 6. The launch vehicleof claim 5, wherein said VG inlet includes an inlet cover.
 7. A methodof propulsion using a ramjet, comprising: providing a high-speed-launchramjet boost (HSLRB) stage including a frame including a front portionand an aft portion, and a fuel tank and fuel pump within said frame anda high-speed-launch ramjet boost (HSLRB) engine within said frameattached to a high-speed launch aircraft which provides a speed of atleast Mach 2.0; said HSLRB engine including: a combustion system forigniting fuel pumped by said fuel pump from said fuel tank, saidcombustion system comprising an igniter, fuel injectors and frameholders; at least one inlet providing a pathway for air to flow withinsaid frame toward said fuel injectors; a variable geometry (VG) nozzlehaving a nozzle actuator at said aft portion for exhausting exhaust gasfrom combustion of said fuel by said combustion system, and a processorcoupled to receive sensing signals from at least one of a pressuresensor and a temperature sensor during flight, wherein said processorprovides control signals to said nozzle actuator for dynamicallycontrolling an aperture size of said VG nozzle, carrying said HSLRBstage to a speed of at least Mach 1.5 during flight of said high-speedlaunch aircraft; igniting said HSLRB engine while attached to saidhigh-speed launch aircraft when at a speed of at least 2.0 Mach, andseparating said HSLRB stage from said high-speed launch aircraft aftersaid igniting.
 8. The method of claim 7, wherein said HSLRB enginegenerates sufficient excess thrust to separate from said high-speedlaunch aircraft and accelerate to a speed of at least 3 Mach morerelative to its speed at a time of said separating.
 9. The method ofclaim 7, wherein said inlet comprises a VG inlet having an inletactuator, and wherein said processor provides control signals to saidinlet actuator for dynamically controlling a geometry of said VG inlet.10. The method of claim 9, wherein said VG inlet includes a frangibleinlet cover, wherein said frangible inlet cover is shattered before saidigniting.
 11. A high-speed-launch ramjet boost (HSLRB) engine,comprising: A combustion system for igniting fuel pumped by a fuel pumpfrom a fuel tank, said combustion system comprising an igniter, fuelinjectors and frame holders; at least one variable geometry (VG) inlethaving an inlet actuator providing a pathway for air to flow toward saidfuel injectors; a variable geometry (VG) nozzle having a nozzle actuatorfor exhausting exhaust gas from combustion of said fuel by saidcombustion system, and a processor coupled to receive sensing signalsfrom at least one of a pressure sensor and a temperature sensor duringflight, wherein said processor provides control signals to said nozzleactuator for dynamically controlling an aperture size of said VG nozzleand said processor provides control signals to said inlet actuator fordynamically controlling a geometry of said VG inlet.
 12. The HSLRBengine of claim 11, wherein said VG inlet includes an inlet cover.